A lift generating blade has been a key component in aircraft propulsion since the inception of powered flight. As with most aircraft structures, performance and structural margins are thin, and therefore blade structures are usually designed to provide the highest practical strength-to-weight and stiffness-to-weight ratios within bounding airfoil profiles. Blade structures, chosen for efficiency in blade bending, commonly take the form of a simple I-beam or box beam centered at the thickest portion of the blade airfoil profile. In some more advanced designs the box beam may be multi-celled to help prevent buckling while allowing for minimum thickness box upper and lower surfaces (often referred to as the beam caps). In addition to the demanding requirements on the blade to act as an efficient beam in bending, many rotor and propeller blades must rotate about their long axis (termed blade feathering axis or blade pitch axis), to provide variable thrust or lift. This necessitates that the airfoil shape of the blade transition to a circular shape or that the blade be mechanically fastened to a rotating member. This circular beam root shape is less effective than a comparably sized (in terms of area) box beam or I-beam in bending. The inefficiency of the circular shape is compounded by the characteristic of cantilevered beams to have a maximum moment at the root of the beam, thus making the beam root region the most important for beam stiffness.
Across the spectrum of aircraft applications for lift-generating blade structures there are a wide variety of configurations that create different demands on the blade structure. Although strength, weight and stiffness are important for most blade structures, several other factors are involved that are application specific and vary between aircraft configurations. A blade structure in a vertical lift application such as that found in a conventional helicopter rotor has very different constraints compared to those of a horizontal flight vehicle such as a propeller aircraft. 
In the case of a conventional helicopter rotor blade, the root of the blade is often hinged, and thus there are little or no bending moments present at the root of the blade structure. The blade is kept radially extended and somewhat stiffened by the centrifugal force as the blade spins around the center axis. This centrifugal force helps to relieve bending in the blade beam and the presence of a hinge at the root eliminates the build up of bending loads from the tip to the root of the blade. The airflow through the rotor disk in a hovering helicopter (commonly termed axial flow) and the edgewise airflow in helicopter forward flight near the rotor axis of rotation are relatively slow, and thus the aerodynamic shape of the blade of a conventional helicopter near the root is not critical to rotor performance.
A typical prior art helicopter rotor blade root and rotor hub attachment can be seen in FIG. 1, adapted from U.S. Pat. No. 4,251,309 to Class, et al. This and all other extrinsic materials discussed herein are incorporated by reference in their entirety. Where a definition or use of a term in an incorporated reference is inconsistent or contrary to the definition of that term provided herein, the definition of that term provided herein applies and the definition of that term in the reference does not apply. In FIG. 1 the blade root 100 comprises an inboard blade aerodynamic portion 120, an unfaired structural root portion 110, and a hub attachment 130. The rectangular box shape in the blade structure 110 continues through to a pinned or otherwise mechanically fastened joint to the hub assembly, which allows the blade to feather or pitch about the blade long axis. In this configuration both the structural beam bending requirements and the aerodynamic constraints due to axial and edgewise flow are relatively low. Thus it can be seen that the structure is not faired, and the blade beam is not optimized for bending.
In the case of an aircraft propeller; the system of blades is moving in the axial direction at a much greater speed than a helicopter (as high as Mach 0.75). Due to these high axial flow rates, the aerodynamic profile of the blade root is critical to the high speed performance of the propeller system. However, unlike helicopter blades, the propeller blade does not support the weight of the aircraft in vertical flight, and is usually of a much smaller overall diameter. Thus the moment loads at the root of the propeller blade are relatively low, and the blade-to-hub attachment depth can be much thinner. As can be seen in FIG. 2, adapted from U.S. Pat. No. 6,155,784 to Carter, a prior art propeller assembly 200 comprises a blade aerodynamic lifting surface portion 210, a blade structure or spar portion 220, a blade shank and rotation portion 230,  and a spinner fairing portion 240. It can be seen the transition from the lifting surface portion 210 to the blade shank and rotation portion 230, can be rapid and still provide propeller aerodynamic efficiency without undue weight.
FIG. 3, adapted from U.S. Pat. No. 4,810,167 to Spoltman, et al., shows a prior art propfan rotor propulsion system 300. Such a system is designed for very high axial speeds of Mach 0.7 to 0.8, and a resulting blade 310 is very thin even at the root section, where it abruptly terminates at a circular turntable 320, used for feathering the blade 310. Such a thin blade has a low capacity for carrying root bending moments.
In the case of a tilt-rotor aircraft both hover and forward flight regimes are possible. When a tiltrotor aircraft operates in airplane cruise mode, the propeller or proprotor operates in substantially high axial flow, (above Mach 0.45). A tiltrotor aircraft usually also has the capability of operating as a helicopter, where axial flow rates are low. Known prior art flying tiltrotor aircraft have what is termed a gimbaled rotor, which allows the rotor to pivot about its center, thus substantially preventing the transfer of high blade root moment loads to the fixed-system hub or mast. In such assemblies, the blade root is typically kept faired close to the hub fairing so that in forward flight the propeller is aerodynamically efficient. However, because moment loads are low in the gimbaled rotor blade, the root structural shape is usually a round spar, structurally inefficient, but simpler in geometry and manufacture for the transition from blade lifting surface portion to blade shank and rotation mechanism.
Hingeless rotor helicopters are known in the prior art. In these designs the rotor blade is cantilevered at the root both in blade bending up-down (termed flap) and in forward-aft (termed lag) while supported on mechanical or elastomeric feather bearings. Such a rotor can be configured to provide rotor control moments for aircraft roll and pitch that are not possible with conventional gimbaled helicopters. In this case, a rotor blade root moment is developed, but as in other helicopters, the axial flow rate is low, and thus again the root aerodynamic profile is not critical to rotor performance. An exemplary hingeless rotor system for a helicopter is that of the Optimum Speed Rotor, U.S. Pat. No. 6,007,298 to Karem and implemented on the Boeing™ Hummingbird A160 helicopter. 
Some advanced technology composite propeller blades have been manufactured using braided pre-forms or composite winding machines. Composite materials can be tailored to place material in directions only as needed to carry prevailing loads on a structure. In the case of a braided pre-form, a number and direction of fibers is determined for the composite structure and a woven “sock” is created, containing only the fibers required for the loads. This sock is placed dry over a mold, and then infused with resin to make the final composite blade part. A similar manufacturing method is composite winding of structures. The prior art all-composite blade of the Airbus™ A400M transport aircraft is manufactured on a winding machine, carrying all the fiber strands (tows in the industry vernacular) and interweaving them on the part essentially simultaneously.
Both woven performs and wound structures have the drawback that the weave by nature creates a slight bend or kink in the fiber as it moves above or below the crossing fiber. These bends reduce the overall strength of the material in tension and compression. A preferred method to reduce this fiber kinking is to create the structure using unidirectional tapes of fibers. Layering these tapes one on top of the other creates an uninterrupted stronger structure, but can be labor intensive to manufacture. Advances in automated fiber placement machines make it possible to wrap unidirectional composite tapes over large complex parts.
The blades of high speed or high efficiency hingeless or rigid rotors or propellers can greatly benefit from achieving high ratios of strength to weight and stiffness to weight within the thin airfoils required for efficiency at substantial axial or edgewise speed. The use of a combination of a rigid rotor and a tiltrotor aircraft as described in U.S. Pat. No. 6,641,365 to Karem provides several advantages. That lightweight hingeless rotor system allows for variable rotor speed, increasing efficiency in both of the two disparate hover and aircraft flight regimes. In both helicopter and airplane modes, a stiff hingeless rotor provides rotor control moments not achievable in the current gimbaled tiltrotor aircraft. However, in this configuration the rotor blade root is subject to two demanding criteria: The blade root must carry large moment loads during hover, as the entire vehicle is lifted by a large diameter unhinged rotor, and in airplane flight mode the blade root must have low aerodynamic drag as the vehicle could see axial flow rates through the rotor disk approaching Mach 0.65. Therefore, it is apparent that a blade structure optimized for high structural root moment capability at low blade root weight and  aerodynamic efficiency at high speed axial flow is still required in the field of hingeless tiltrotor aircraft, and will also substantially benefit hingeless rotor helicopters and compound helicopters with high edgewise speed.
Designs for metal composite blade retention systems have been suggested as in U.S. Pat. No. 3,734,642 to Dixon et al. which suggests one method wherein a composite blade is bonded to an external titanium cuff structure which is in turn bolted to a bearing to allow feathering of the blade. In this design, a backup internal cuff acts as a redundant load path in the case of bond failure on the first cuff to composite interface. A conical flair of the blade root and cuff forces a wedging action between the inner cuff structure and the composite in the event of a primary bond failure. FIG. 4, adapted from the Dixon patent, illustrates a blade root and attachment 400 design comprising a blade spar 410, cuff 430, and attachment to a hub 420. It is important to note that in the Dixon patent and other known prior art designs, the blade spar 420 and cuff 430 both have circular cross sections.